A conventional liquid fuel rocket engine has a propellant (fuel and oxidizer) injector and a nozzle shaped combustion chamber into which the propellant is injected and where its constituents are mixed and oxidized or combusted.
There are a number of proven fuel and oxidizer combinations, each with specific performance characteristics. They differ in the amount of energy released when they combust and in the thermodynamics of their reactions. By the nature of their reactions all the propellant combinations are exothermic, that is, once the fuel and oxidizer are mixed, the reaction is initiated and energy is released. Once the reaction has been initiated, the release of energy from the propellant combination will drive the reaction.
The efficiency of the reaction, that is, the degree to which both propellant components are completely reacted, is largely dependent on the thoroughness of the mixing of the two components. Incomplete mixing results in unreacted fuel or oxidizer and a corresponding loss in efficiency.
The propellant injector of the rocket engine implements the mixing of the fuel and oxidizer components of the propellant. A typical injector can have from several to several thousand orifices through which the fuel and oxidizer are introduced into the combustion chamber. The orifices direct the fuel and the oxidizer so that they form spray fans which mix and commence combustion at preselected points within the combustion chamber, typically immediately down stream of the injector. Numerous orifice arrangements which form fuel only, oxidizer only or bipropellant fans exist.
The combustion of rocket propellant generates gas temperatures which generally exceeds the melting temperature of most known materials used in the construction of chamber walls. Without cooling the chamber walls would deteriorate and ultimately melt.
Further, the heating of the combustion chamber walls can lead to an overheating of the injector as a result of heat transfer between them via their interface. This can result in a vapor lock in the injector, and engine failure.
There are two main approaches to cooling the combustion chamber walls of a rocket engine. Regenerative cooling circulates one or both of the propellant components through the walls of the chamber. The propellant component acts as a coolant and carries away the heat which is eventually returned to the combination gases. The approach has only limited utility. It can not provide sufficient cooling for small engines because their propellant flow is too low and this approach may not be suited for use with large, high-pressure engines for other reasons.
The other often practiced approach employs film cooling in which the orifice pattern of the propellant injector generates two propellant flows, a central core flow and a peripheral or curtain flow which surrounds or envelopes the core flow. In the core flow the propellant is well mixed and combusted in the core which is some distance radially inward of the chamber wall. The cooler curtain is formed by unmixed and therefore uncombusted propellant directed by injector orifices toward the chamber walls. The unmixed propellant forms a cool gas film or curtain over the chamber wall which separates it from the very hot core flow. The film absorbs heat by evaporation of the small fuel and/or oxidizer droplets ejected by the injector and thus insulates the combustion chamber wall from the heat of the core flow. This method can be used with most rocket engines, but the film of uncombined propellant in contact with the combustion chamber wall is disadvantageous.
The propellant film evaporates and decomposes from its exposure to the heat. Decomposition products react with unmixed propellant to create a variety of aggressive chemical species which can chemically react with typical chamber wall materials such as copper, nickel, platinum, iridium, gold, rhenium and columbium. This corrodes and deteriorates the chamber wall and can lead to its failure.
Further, the propellant film can undergo spontaneous thermal decomposition, resulting in transient species which can be the source of additional chemical attack on the combustion chamber wall.
Partially mixed propellant can result in localized concentrations of fuel and oxidizer existing side by side. The boundaries between these concentrations create an environment where a spectrum of combustion chemistry species are generated, including many nonequilibrium species not normally found in other combustion devices. When these species come in contact with the combustion chamber wall, they cause a condition known as a streaking, blanching or scalloping of the walls. They can also attack the injector orifices, resulting in what is commonly called bell-mouthing.
A further disadvantage of prior art film cooling is that it decreases the efficiency of the rocket engine. The quantity of the propellant used for cooling is significant and, to a substantial extent, its use as a coolant causes it to be lost to the system as propellant. This can result in an appreciable loss of efficiency.